The present invention relates to gas turbine engine control systems and more particularly to stall warning systems for such engines.
The phenomenon of compressor stall has become an important limiting factor in the operation of gas turbine engines as their performance characteristics have improved. In modern gas turbine engines, upon acceleration or under high altitude and low airspeed flight conditions, unstable flow may develop in the compressor which can lead to an increase in exhaust gas temperature and mechanical vibration along with a simultaneous reduction in cooling air supplied to the turbine wheel. These conditions describe "compressor stall" and can lead to turbine wheel failure if the compressor stall is not recognized and corrective action not taken. Turbine wheel failure during engine operation can lead to severe engine and aircraft damage.
To avoid such damage, control systems have been proposed to detect compressor stall and either signal the aircraft pilot or automatically compensate so as to prevent further stall. It has been found that functions of certain engine operating parameters can provide an indication of incipient stall conditions. Systems have been designed which by monitoring these parameters can detect compressor stall. Several previous attempts to develop stall warning indicators have used either engine operating pressures or pressure ratios as input parameters. However, these systems have been unsuccessful because no dependable criteria on pressure can be established which present a reliable indication of compressor stall alone. For example, transient pressures or pressure ratios, which normally occur during safe engine operation, can provide a false indication of compressor stall. Further, these prior art attempts have required expensive aircraft modification prior to their installation.
Compressor stalls can also be evidenced by certain engine rotor speed, sonic amplitude, or gas temperature relationships. Although a tachometer provides the most positive indication of impending stall it is not sufficiently responsive, for by the time the engine speed can visibly be seen to decrease, stall has already begun and it is too late to react and prevent turbine shut down. Compressor sonic amplitude is not in itself necessarily indicative of the compressor operating condition. Variations in engine mechanical tolerance from different manufacturers and metal spurs produced by localized metal fatigue can result in widely differing sonic amplitudes and confused sonic signals.
The most practical compressor stall detection systems have employed combinations of turbine inlet temperature and compressor rotor rotational speed to detect incipient engine stall. Several systems have been proposed which utilize various combinations of the above mentioned parameters including differences, square roots, and dividends to indicate stall. However, these systems, as previously proposed, have been deficient in sensitivity and lacked ability to operate under all flight conditions.